1. Field of the Invention
The present invention relates generally to spacecraft and more particularly to spacecraft radiation sensors.
2. Description of the Related Art
FIG. 1 illustrates an exemplary body-stabilized satellite 10 which travels along an orbit path 12 that defines an orbit plane 13 about the earth 15. Carried on or within the satellite's body 16 are various operational systems, e.g., a communications system which includes antennas 18, a propulsion system which includes thrusters 20 and an energy-generation system which includes solar wings 22 and 24.
Typically, a spacecraft must control its attitude in order to carry out system functions for which it was designed. In the case of the satellite 10, for example, the attitude of the satellite's body must be controlled to direct the antennas 18 towards the earth 15 and to arrange the solar wings 22 and 24 so that they can rotate daily and receive solar energy arriving along a sun line 26. In order to appropriately alter the satellite's attitude, (e.g., with on-board momentum wheels and/or the thrusters 20), its attitude must first be sensed and determined.
Satellite attitude is typically defined as the relationship between a coordinate frame of the satellite's body and an external coordinate frame such as the orthogonal orbit normal frame 30 of FIG. 1 which has a yaw axis 31 directed at the center of the earth 15, a pitch axis 32 that is orthogonal to the orbit plane 13 and a roll axis 33 in the direction of motion.
To find the attitude of a spacecraft, the directions of various radiation sources such as celestial sources (e.g., the earth, the sun and stars) and artificial sources (e.g., a radar beacon signal from an earth-based antenna) are sensed and the satellite's inertial attitude obtained relative to these radiaiton sources.
For example, an earth sensor can determine a nadir vector between a satellite and the earth's center. Knowledge of the nadir vector can be used to orient the roll-pitch plane of the satellite 10 of FIG. 1. However, because the nadir vector provides no attitude information relative to rotation about the yaw axis 31, an earth sensor is typically supplemented with a sun sensor which can determine the direction of the sun line 26 of FIG. 1.
Conventional sun sensors, however, typically have a limited field-of-view. Accordingly, sun sensors can only track the sun over a limited portion of an earth satellite's orbit and other attitude sensors (e.g., gyroscopes) are typically used for attitude determination over the remainder of the orbit. Because gyroscopes have a rate drift, more accurate attitude information from celestial radiation sensors (e.g., sun sensors) is generally used to calibrate them. If wider field-of-view radiation sensors were available, the gyroscopes could be calibrated over a greater portion of a satellite's orbit and relied upon over a smaller portion of the orbit (i.e., over a reduced time period). Because of both of these effects, the accuracy of satellite attitude determination would be enhanced.
Sun sensors have been described in various references (e.g., see Morgan, Walter L., et al., Communications Satellite Handbook, John Wiley & Sons, New York, 1989, pp. 648-649). An exemplary sun sensor structure 40 (e.g., see U.S. Pat. No. 4,999,483) is shown in FIG. 2A. The sun sensor 40 has a detector array 42 positioned within a sensor body 43. The detector array is formed with a plurality of detector elements 44. The body includes a shield member 45 which forms a radiation aperture 46.
With the sensor 40 in a first attitude with respect to the sun, a first radiation sheet 48A passes through the radiation aperture 46 and illuminates a detector element 44A in the detector array 42. When sensor 42 has a different second attitude with respect to the sun, a different second radiation sheet 48B is received through the radiation aperture 46 at a different angle and it illuminates a different detector element 44B in the detector array 42 (detailed optical effects (e.g., diffraction) are ignored in this brief description of radiation passage through the radiation aperture).
Along a sensor axis 50, therefore, illuminated detector elements are indicative of attitudes between the sun sensor 40 and the sun. For example, successive illumination of detector elements 44A and 44B indicates that attitude has changed over a sensor angle 51. Along an aperture axis 52, radiation is received over an aperture angle 53 but the sun's position along this axis is indeterminate.
FIG. 2B is similar to FIG. 2A with like elements indicated by like reference numbers. FIG. 2B illustrates a radiation solid angle 58 which is received by the sun sensor 40 and which defines its field-of-view. The solid angle 58 includes a detector angle 60 along the sensor axis 50 and an aperture angle 62 along the aperture axis 52. In a specific sun sensor structure, the magnitudes of the detector angle 60 and the aperture angle 62 are functions of the spatial size of the sensor array 42 and the radiation aperture 46 and the spatial arrangement of these elements.
In FIG. 2B, the solid angle 58 has an exemplary solid-angle cross section 70 which is redrawn in FIG. 2C along with the sensor axis 50 and the aperture axis 52. For illustrative purposes, a series of sensor lines 72 are positioned across the solid-angle cross section 70. These sensor lines are orthogonal to the sensor axis 50 and they each correspond to a respective one of the detector elements (44 in FIG. 2A).
The spacing between the sensor lines 72 is a function of the spatial layout of the sun sensor (40 in FIG. 1A) and indicates the resolution of the sensor along the sensor axis 50. The horizontal extent of the sensor lines 72 is indicative of the field-of-view (at a respective solid-angle cross section) along the aperture axis 52. These sensor lines can be considered to define respective sensor planes in the sensor solid angle 58 of FIG. 2B. All of these sensor planes are orthogonal to the sensor axis 50.
Various sun sensor modifications have been proposed. In one conventional type of sun sensor, for example, each detector element (44 in FIG. 2A) is formed with a row of elements which are coded by a mask so that, when illuminated, they generate a digital code (e.g., a gray code) that is indicative of that respective attitude between the sensor and the sun.
Two-axis sun sensors are typically formed by combining a pair of single-axis sun sensors. For example, the two-axis sun sensor 80 of FIG. 2D includes two single-axis sun sensors 40A and 40B which have been positioned in proximity to each other and with their sensor axes 50A and 50B in an orthogonal relationship. Accordingly, their sensor apertures 46A and 46B are also in an orthogonal relationship.
FIG. 2E shows a solid-angle cross section 90 for the two-axis sun sensor 80. This radiation cross section is similar to the cross section 70 of FIG. 2C in that it has a series of sensor lines 92A which are orthogonal to the sensor axis 50A and which correspond to sensor elements of the single-axis sun sensor 40A. In addition, it has a series of sensor lines 92B which are orthogonal to the sensor axis 50B and which correspond to sensor elements of the single-axis sun sensor 40B and which are orthogonal to the sensor lines 92A. Exemplary sensor lines 94 and 96 now uniquely determine an attitude point 97 (i.e., an attitude vector) in the radiation solid angle that is received by the two-axis sun sensor 90.
As mentioned above, however, conventional sun sensors have a limited field-of-view which degrades attitude determination because it increases the time over which attitude must be determined by less accurate means (e.g., gyroscopes). The limited field-of-view is apparent in FIG. 2E from an exemplary sun track 98 which enters and leaves the sensor's solid angle at attitude points 99E and 99L.